Gas turbine engine with blade channel variations

ABSTRACT

A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor including a rotor hub and an array of blades circumferentially spaced about the rotor hub, a geared architecture, a compressor section and a turbine section. Each blade includes pressure and suction sides and extends in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip, adjacent blades in the array of blades including a first blade and a second blade, a facing pressure side of the first blade and a facing suction side of the second blade defining a channel having a width that varies in a chordwise direction between the facing pressure and suction sides at a given span position of the first and second blades. The width at each pressure side location of the first blade along the channel is defined as a minimum distance from the respective pressure side location to a location along the suction side of the second blade, and the width of the channel converges in the chordwise direction to establish a throat.

CROSS-REFERENCE TO RELATED APPLICATION

The present disclosure is a continuation of U.S. patent application Ser.No. 16/519,475, filed Jul. 23, 2019, which is a continuation of U.S.patent application Ser. No. 15/045,343, filed Feb. 17, 2016, which is acontinuation of U.S. patent application Ser. No. 14/699,322, filed Apr.29, 2015 and issued as U.S. Pat. No. 9,470,093, which claims the benefitof U.S. Provisional Patent Application No. 62/134,760, filed Mar. 18,2015.

BACKGROUND

This disclosure relates generally to a fan stage for gas turbineengines, and more particularly to a relationship between channel widthrelative to span for adjacent pairs of fan blades and correspondingperformance and stall margin characteristics.

A turbine engine such as a gas turbine engine typically includes a fansection, a compressor section, a combustor section and a turbinesection. The fan section includes a plurality of fan blades spacedcircumferentially to define a plurality of channels. The fan bladescompress a portion of incoming air through the channels to producethrust and also deliver a portion of air to the compressor section. Airentering the compressor section is compressed and delivered into thecombustor section where it is mixed with fuel and ignited to generate ahigh-speed exhaust gas flow. The high-speed exhaust gas flow expandsthrough the turbine section to drive the compressor and the fan section.The compressor section typically includes low and high pressurecompressors, and the turbine section includes low and high pressureturbines.

The efficiency of a gas turbine engine depends on many differentfactors. The fast-moving air creates flow discontinuities or shocks thatresult in irreversible losses. In addition to contributing to theoverall efficiency of the engine, the fan module contributessignificantly to the weight of the engine. As such, features that reducethe collective weight of the fan blades or the weight of the module ingeneral contribute positively to aircraft fuel consumption.

SUMMARY

A fan section for a gas turbine engine according to an example of thepresent disclosure includes a rotor hub defining an axis, and an arrayof airfoils circumferentially spaced about the rotor hub. Each of theairfoils include pressure and suction sides between a leading edge and atrailing edge and extending in a radial direction from a 0% spanposition at an inner flow path location to a 100% span position at anairfoil tip, facing pressure and suction sides of adjacent airfoilsdefining a channel in a chordwise direction having a width between thefacing pressure and suction sides at a given span position of theadjacent airfoils. The width at each pressure side location along thechannel is defined as a minimum distance to a location along the suctionside. The width diverges without converging along the channel for atleast some of the span positions, and the width converging and divergingalong the channel for at least some span positions greater than 5% spanand less than half of the span positions. Each of the array of airfoilshas a solidity defined by a ratio of an airfoil chord over acircumferential pitch. The solidity is between about 1.6 and about 2.5for each of the span positions in which the width converges along thechannel.

In a further embodiment of any of the foregoing embodiments, the widthdiverges without converging along the channel at span positions greaterthan about 16% span.

In a further embodiment of any of the foregoing embodiments, the widthconverges and diverges along the channel at span positions greater thanor equal to about 10% span.

In a further embodiment of any of the foregoing embodiments, the widthdiverges without converging at span positions from 100% span to lessthan or equal to 90% span.

In a further embodiment of any of the foregoing embodiments, thesolidity at the tip of each of the array of airfoils is less than orequal to about 1.2.

In a further embodiment of any of the foregoing embodiments, the widthconverges and diverges for less than or equal to about 20% of the spanpositions.

In a further embodiment of any of the foregoing embodiments, a ratio ofthe width to the solidity at each span position is greater than or equalto about 0.50.

In a further embodiment of any of the foregoing embodiments, thesolidity at each span position is greater than or equal to about 0.8.

In a further embodiment of any of the foregoing embodiments, the arrayof airfoils includes 20 or fewer airfoils.

In a further embodiment of any of the foregoing embodiments, flowthrough the channel at span positions where the width converges anddiverges along the channel corresponds to a leading edge relative machnumber less than or equal to about 0.8 Mach at cruise.

In a further embodiment of any of the foregoing embodiments, a staggerangle of each of the array of airfoils relative to the axis is less thanor equal to about 16 degrees at each of the span positions in which thewidth converges and diverges along the channel.

In a further embodiment of any of the foregoing embodiments, the widthconverges and diverges for less than or equal to about 20% of the spanpositions, and the width diverges without converging at span positionsfrom 100% span to less than or equal to about 80% span.

In a further embodiment of any of the foregoing embodiments, the widthconverges along the channel at a location spaced a distance from aninlet of the channel, the distance being greater than a radius definedby the leading edge at the same span position.

In a further embodiment of any of the foregoing embodiments, thesolidity at the 0% span position is greater than or equal to about 2.3.

A gas turbine engine according to an example of the present disclosureincludes a combustor section arranged between a compressor section and aturbine section. A fan section has a rotor hub and an array of airfoilscircumferentially spaced about the rotor hub to define a plurality ofchannels. Each of the array of airfoils includes pressure and suctionsides and extend in a radial direction from a 0% span position at aninner flow path location to a 100% span position at an airfoil tip,facing pressure and suction sides of adjacent airfoils defining achannel in a chordwise direction having a width between the facingpressure and suction sides at a given span position of the adjacentairfoils. The width at each pressure side location along the channel isdefined as a minimum distance to a location along the suction side. Thewidth converges along the channel for at least some span positionsgreater than 5% span and less than half of the span positions. A staggerangle of each of the array of airfoils relative to the axis is less thanor equal to about 16 degrees at span positions converging and divergingalong the channel.

In a further embodiment of any of the foregoing embodiments, the widthconverges and diverges for less than or equal to about 20% of the spanpositions.

In a further embodiment of any of the foregoing embodiments, the widthdiverges without converging at span positions from 100% span to lessthan or equal to 90% span.

In a further embodiment of any of the foregoing embodiments, each of thearray of airfoils has a solidity defined by a ratio of an airfoil chordover a circumferential pitch, and a ratio of the width to the solidityat each span position is greater than or equal to about 0.50.

In a further embodiment of any of the foregoing embodiments, thesolidity at each span position is greater than or equal to about 0.8.

In a further embodiment of any of the foregoing embodiments, the arrayof airfoils includes 20 or fewer airfoils.

These and other features of this disclosure will be better understoodupon reading the following specification and drawings, the following ofwhich is a brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a gas turbine engine.

FIG. 2A is a perspective view of a fan section of the engine of FIG. 1.

FIG. 2B is a schematic cross-sectional view of the fan section of FIG.2A.

FIG. 2C is a schematic view of a cross-section of an airfoil of FIG. 2Bsectioned at a particular span position and depicting directionalindicators.

FIG. 3A is a schematic view of airfoil span positions.

FIG. 3B is a perspective view of sections of the airfoil of FIG. 2A atvarious span positions.

FIG. 3C is a schematic representation of a dihedral angle for anairfoil.

FIG. 4 is a schematic view of adjacent airfoils depicting a leading edgegap, or circumferential pitch, and a chord of the airfoil.

FIG. 5 graphically depicts curves for several example airfoil chord/gapratios relative to span, including one prior art curve and severalinventive curves according to this disclosure.

FIG. 6A is a schematic view of the adjacent airfoils depicting pointpairs indicating widths along a channel between the adjacent airfoils ata first span position.

FIG. 6B graphically depicts curves for example channel widths betweenthe adjacent airfoils of FIG. 6A normalized by trailing edge pitch.

FIG. 6C is a schematic view of the adjacent airfoils of FIG. 6Adepicting point pairs indicating widths along a channel between theadjacent airfoils at a second span position.

FIG. 6D graphically depicts curves for example channel widths betweenthe adjacent airfoils of FIG. 6C normalized by trailing edge pitch.

FIG. 6E is a schematic view of the adjacent airfoils of FIG. 6Adepicting point pairs indicating widths along a channel between theadjacent airfoils at a third span position.

FIG. 6F graphically depicts curves for example channel widths betweenthe adjacent airfoils of FIG. 6E normalized by trailing edge pitch.

FIG. 7 graphically depicts curves for several example relative Machnumbers relative to span, including one prior art curve and severalinventive curves according to this disclosure.

FIG. 8 graphically depicts curves for several example airfoil staggerangles (α) relative to span, including one prior art curve and severalinventive curves according to this disclosure.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a star gearsystem, a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3 and the low pressure turbine46 has a pressure ratio that is greater than about five. In onedisclosed embodiment, the engine 20 bypass ratio is greater than aboutten (10:1), the fan diameter is significantly larger than that of thelow pressure compressor 44, and the low pressure turbine 46 has apressure ratio that is greater than about five 5:1. The engine 20 in oneexample is a high-bypass geared aircraft engine. In another example, theengine 20 bypass ratio is greater than about twelve (12), the gearedarchitecture 48 has a gear reduction ratio of greater than about 2.6 andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five. Low pressure turbine 46 pressure ratio is pressure measuredprior to inlet of low pressure turbine 46 as related to the pressure atthe outlet of the low pressure turbine 46 prior to an exhaust nozzle.The geared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.3:1. It should be understood, however, that theabove parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present disclosure is applicable toother gas turbine engines including direct drive or non-gearedturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than or equal to about1.50, with an example embodiment being less than or equal to about 1.45.“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram °R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1200ft/second, or more narrowly less than about 1150 ft/second.

In one example, the fan section 22 includes a hardwall containmentsystem 23 arranged about the engine axis A and spaced radially from thefan blades 42. The hardwall containment system 23 is configured tocontain, and absorb the impact of, a fan blade 42 separating from a fanhub 76 or a fragment thereof. In some embodiments, the hardwallcontainment system 23 is a hard ballistic liner applied to the nacelleor fan case 15. The hard ballistic liner can include a rigid materialsuch as a resin impregnated fiber structure, metallic structures, orceramic structures.

Various materials and structures of the fan case 15 and/or hardwallcontainment system 23 are contemplated. In some embodiments, the fansection 22 includes a composite fan case 15 made of an organic matrixcomposite. The organic matrix composite can include a matrix materialand reinforcement fibers distributed through the matrix material. Thereinforcement fibers may be discontinuous or continuous, depending uponthe desired properties of the organic matrix composite, for example. Thematrix material may be a thermoset polymer or a thermoplastic polymer.The reinforcement fibers may include carbon graphite, silica glass,silicon carbide, or ceramic. Given this description, one of ordinaryskill in the art will recognize that other types of matrix materials andreinforcement fibers may be used. The disclosed arrangements of thecomposite fan case 15 reduce the overall weight of the engine, therebyimproving aircraft fuel consumption.

Referring to FIGS. 2A-2C, the fan 42 includes a rotor 70 having an arrayor row 72 of airfoils or blades 74 that extend circumferentially aroundand are supported by the fan hub 76. Any suitable number of fan blades74 may be used in a given application. The hub 76 is rotatable about theengine axis A. The array 72 of fan blades 74 are positioned about theaxis A in a circumferential or tangential direction Y. Each of theblades 74 includes an airfoil body that extends in a radial spandirection R from the hub 76 between a root 78 and a tip 80, in a chorddirection H (axially and circumferentially) between a leading edge 82and a trailing edge 84 and in a thickness direction T between a pressureside P and a suction side S.

Each blade 74 has an exterior surface 88 providing a contour thatextends from the leading edge 82 aftward in a chord-wise direction H tothe trailing edge 84. The exterior surface 88 of the fan blade 74generates lift based upon its geometry and directs flow along the coreflow path C and bypass flow path B. The fan blade 74 may be constructedfrom a composite material, or an aluminum alloy or titanium alloy, or acombination of one or more of these. Abrasion-resistant coatings orother protective coatings may be applied to the fan blade 74.

A chord, represented by chord dimension (CD), is a straight line thatextends between the leading edge 82 and the trailing edge 84 of theblade 74. The chord dimension (CD) may vary along the span of the blade74. The row 72 of blades 74 also defines a circumferential pitch (CP)that is equivalent to the arc distance between the leading edges 82 ortrailing edges 84 of neighboring blades 74 for a corresponding spanposition. The root 78 is received in a correspondingly shaped slot inthe hub 76. The blade 74 extends radially outward of a platform 79,which provides the inner flow path. The platform 79 may be integral withthe blade 74 or separately secured to the hub 76, for example. A spinner85 is supported relative to the hub 76 to provide an aerodynamic innerflow path into the fan section 22.

Referring to FIGS. 3A-3B, span positions are schematically illustratedfrom 0% to 100% in 10% increments to define a plurality of sections 81.Each section at a given span position is provided by a conical cut thatcorresponds to the shape of segments the bypass flowpath B or the coreflow path C, as shown by the large dashed lines (shown in FIG. 3A). Inthe case of a fan blade 74 with an integral platform, the 0% spanposition corresponds to the radially innermost location where theairfoil meets the fillet joining the airfoil to the platform 79. In thecase of a fan blade 74 without an integral platform, the 0% spanposition corresponds to the radially innermost location where thediscrete platform 79 meets the exterior surface of the airfoil (shown inFIG. 2B). A 100% span position corresponds to a section of the blade 74at the tip 80.

In some examples, each of the blades 74 defines a non-linear stackingaxis 83 (shown in FIG. 3B) in the radial direction R between the tip 80and the inner flow path location or platform 79. For the purposes ofthis disclosure, “stacking axis” refers to a line connecting the centersof gravity of airfoil sections 81. In some examples, each fan blade 74is specifically twisted about a spanwise axis in the radial direction Rwith a corresponding stagger angle α at each span position and isdefined with specific sweep and/or dihedral angles along the airfoil 74.Airfoil geometric shapes, stacking offsets, chord profiles, staggerangles, sweep and dihedral angles, among other associated features, canbe incorporated individually or collectively to improve characteristicssuch as aerodynamic efficiency, structural integrity, and vibrationmitigation, for example.

In some examples, the airfoil 74 defines an aerodynamic dihedral angle D(simply referred to as “dihedral”) as schematically illustrated in FIG.3C. An axisymmetric stream surface S passes through the airfoil 74 at alocation that corresponds to a span location (FIG. 3A). For the sake ofsimplicity, the dihedral D relates to the angle at which a line L alongthe leading or trailing edge tilts with respect to the stream surface S.A plane P is normal to the line L and forms an angle with the tangentialdirection Y, providing the dihedral D. A positive dihedral D correspondsto the line tilting toward the suction side (suction side-leaning), anda negative dihedral D corresponds to the line tilting toward thepressure side (pressure side-leaning).

FIG. 4 shows an isolated view of a pair of adjacent airfoils designatedas leading airfoil 74A and following airfoil 74B. As shown, each airfoil74A, 74B is sectioned at a radial position between the root 78 and thetip 80. Chord CD, which is the length between the leading and trailingedges 82, 84, forms an angle, or stagger angle α, relative to theX-direction or plane parallel to the engine's central longitudinal axisA. The stagger angle α can vary along the span to define a twist.

As shown, each airfoil 74 has an asymmetrical cross-sectional profilecharacterized by a mean camber line 75 bisecting a thickness of theairfoil 74 in the chord-wise direction H and a camber angle 77 definedby a projection of the leading and trailing edges 82, 84. The camberangle 77 can differ at various span positions.

The leading edges 82, or trailing edges 84, of the adjacent airfoils 74are separated by gap or circumferential pitch (CP) in the Y-direction,which is a function of blade count. A ratio of chord to gap, which isreferred to as solidity, varies with position along the span.

The geared architecture 48 of the disclosed example permits the fan 42to be driven by the low pressure turbine 46 through the low speed spool30 at a lower angular speed than the low pressure turbine 46, whichenables the low pressure compressor 44 to rotate at higher, more usefulspeeds. The solidity along the span of the airfoils 74 providesnecessary fan operation in cruise at lower speeds enabled by the gearedarchitecture 48, to enhance aerodynamic functionality and efficiency.

The airfoil 74 has a relationship between a chord/gap ratio or solidity(CD/CP) and span position. FIG. 5 illustrates the relationship betweenthe chord/gap ratio (CD/CP) and the average span (AVERAGE SPAN %), whichis the average of the radial position at the leading and trailing edges82, 84. One prior art curve (“PRIOR ART”) is illustrated as well asseveral example curves 92, 93, 94 corresponding to different fanarrangements. The prior art curve corresponds to a fan section having 24blades, characterized by a relatively high solidity along the span.

Curves 92, 93, 94 correspond to fan sections with relatively lowsolidity. Curve 92 is characterized by a chord/gap ratio in a range ofless than or equal to about 2.4 at 0% span to a chord/gap ratio greaterthan or equal to about 1.0 at 100% span. The example curve 92corresponds to a fan section having 18 or fewer blades. Curve 93 ischaracterized by a chord/gap ratio in a range of less than or equal toabout 2.5 at 0% span to a chord/gap ratio greater than or equal to about1.0 at 100% span. The example curve 93 corresponds to a fan sectionhaving 20 or fewer blades. Curve 94 is characterized by a chord/gapratio in a range of less than or equal to about 2.5 at 0% span to achord/gap ratio greater than or equal to about 1.1 at 100% span. Theexample curve 94 corresponds to a fan section having 18 or fewer blades.In some examples, the fan section 22 has 20 or fewer fan blades, morenarrowly 18 or fewer fan blades, or between 18 and 20 fan blades,utilizing any of the techniques of this disclosure.

As shown, the example curves 92, 93, 94 have a lower solidity orchord/gap ratio than the prior art curve. The chord/gap ratio of theinventive curves is less than the prior art curve due in part to a lowerrelative speed of the fan section than the low pressure turbine and arelatively lower pressure ratio, which is enhanced by the gearedarchitecture 48. Low solidity fan blades, including the example fanblade arrangements illustrated by curves 92, 93, 94 corresponding toblade 74, improve the weight of the engine 20, thereby reducing fuelconsumption.

Other chord/gap ratios are contemplated with the teachings of thisdisclosure. In some examples, the chord/gap ratio is less than or equalto about 2.5 along the average span. In another example, the chord/gapratio is greater than or equal to about 1.0 along the average span. Inone example, the chord/gap ratio is less than or equal to about 2.5, andgreater than or equal to about 1.0, along the average span. In otherexamples, the chord/gap ratio is less than or equal to about 1.3, orless than or equal to about 1.1. In some examples, the chord/gap ratiois less than or equal to about 1.0 for at least some of the averagespan.

Fluid dynamic interaction between the fan blades 74 and an incoming airstream generally causes aerodynamic losses such as from shocks orviscous effects. The performance and stability characteristics of thefan section 22 depend on several factors, including the geometry andspatial arrangement of adjacent pairs of fan blades 74. Referring toFIG. 6A, a leading fan blade 74A and a following fan blade 74B arespaced apart in the circumferential direction Y to define a channel 96extending generally in a chordwise direction from a leading edge 82 offan blade 74B. Pressure side P of fan blade 74B and suction side S sideof fan blade 74A define a plurality of segments or channel widths (O) atlocations 98 along the channel 96. For the purposes of this disclosure,each channel width 98 along the channel 96 is a straight-line connectedpoint pair extending from a point 101B on the pressure side P of the fanblade 74B to a point 101A on the suction side S of the fan blade 74Athat is a minimum distance from the point on the pressure side P, andare normalized by trailing edge pitch. In this manner, a connected pointpair can be defined at each location along the channel 96. The channelwidths 98 vary generally in the axial direction X due to contouring ofexterior surfaces 88. The channel widths 98 may vary along the span ofthe blade 74.

The channel 96 is provided with an inlet 99 at the leading edge 82 offan blade 74B and an outlet 100 downstream of the inlet 99. In someexamples, the width of the channel 96 diverges without converging in achordwise direction along the channel 96 for each of the span positions.This arrangement is shown in FIG. 6E at a given span position.

In other examples, at some span positions the channel width 98 convergesand diverges along the channel 96 to define a converging-diverging (C-D)diffuser 101 (shown in FIG. 6A), where 102 is labeled at a throat of theC-D diffuser 101. The throat 102 defines a minimum channel width 98along the channel 96. The throat 102 is located downstream of a positionalong the pressure side P of fan blade 74B at a radius defined by theleading edge 82. Rather, the inlet 99 is characterized in part by thegeometry of the leading edge 82 (best seen in FIG. 6D), whereas thethroat 102 is characterized by the contouring of the pressure side Pdownstream of the inlet 99 (best seen in FIG. 6A). In other examples,the channel 96 has substantially the same minimum channel width 98 fromthe throat 102 to the outlet 100, sometimes referred to as a“converging-straight” configuration.

Various arrangements for the C-D-diffuser 101 are contemplated. The C-Ddiffuser 101 can extend radially from about a 0% span position to spanpositions greater than 5% span, greater than 10% span, greater than 16%span, or greater than about 20% span. In some examples, the channelwidth 98 diverges without converging at 100% span, more narrowly from100% span to about 80% span, from 100% span to about 70% span, or from100% span to about 50% span.

The C-D diffuser 101 can be defined along a range of span positions. Insome examples, the C-D diffuser 101 is defined for about 75% or fewer ofthe span positions, more narrowly about 50% or fewer of the spanpositions, or more narrowly about 30% of the span positions. In otherexamples, the C-D diffuser 101 is defined for about 20% or fewer of thespan positions, more narrowly about 15% or fewer of the span positions,or even more narrowly about 10% or fewer of the span positions.

The C-D diffuser 101 may be configured with various solidity orchord/gap ratios, including any of the chord/gap ratios shown in FIG. 5.In one example, the chord/gap ratio is less than or equal to about 2.2for at least some span positions of the C-D diffuser 101. In oneexample, the chord/gap ratio is less than or equal to about 2.1, or lessthan or equal to about 2.0, for at least some span positions of the C-Ddiffuser 101. In another example, the chord/gap ratio is less than orequal to about 1.5 for at least some span positions of the C-D diffuser101. In other examples, the chord/gap ratio is between about 1.6 andabout 2.5, or more narrowly between about 2.0 and about 2.5, or evenmore narrowly between about 2.3 and about 2.5 at each span position ofthe C-D diffuser 101.

FIG. 6B illustrates example plots 103A, 103B for a ratio between channelwidths (O) and a dimension (tau) of the trailing edge circumferentialpitch CP corresponding to fan blades 74A, 74B of FIG. 6A at 0% spanposition. The channel widths (O) correspond to positions along thechannel 96, including locations 98, 99, 100 and 102. As shown, the ratio(O/tau) decreases from the inlet 99 to throat 102 with respect to theengine axis A (x-axis) and thereafter increases downstream of the throat102 to the outlet 100 to define inflections 105A, 105B. Rather, theratio (O/tau) at the throat 102 is the minimum value for the channelwidths 98 along the channel 96. The C-D diffuser 101 is defined for lessthan half of the span of the channel 96, such that the channel width 98increases or is divergent from the inlet 99 to the outlet 100 for otherportions of the span, and with the ratio (O/tau) generally increasingfrom the inlet 99 to the outlet 100 of the channel 96 (illustrated byFIGS. 6C-6D at about 20% span, for example). Other portions of thechannel 96 are divergent or increase from the inlet 99 to the outlet 100of the channel, as illustrated by FIGS. 6E-6F at about 100% span,according to an example.

Various ratios (O/tau) are contemplated in combination with the channelarrangements of this disclosure. In some examples, the ratio (O/tau) ateach span position is greater than or equal to about 0.50. In oneexample, the ratio (O/tau) at each span position is greater than orequal to about 0.60, and or greater than or equal to about 0.70. Inanother example, the ratio (O/tau) of the throat 102 at each spanposition of the C-D diffuser 101 is between about 0.50 and about 0.85,between about 0.73 and about 0.81, or between about 0.75 and about 0.82.In some examples, the ratio (O/tau) of the throat 102 at each spanposition of the C-D diffuser 101 is greater than about 0.75, or greaterthan or equal to about 0.8. In an example, the ratio (O/tau) of thethroat 102 at each span position of the C-D diffuser 101 is betweenabout 0.79 and about 0.83. These various arrangements further improvefluid stability through the channel 96 and improve fan blade efficiencyin accordance with the teachings of this disclosure.

FIG. 7 illustrates the relationship between the Mach number of the flow(in the rotating frame of reference) through the channel 96 and theblade leading edge span (LEADING EDGE SPAN %) for various arrangementscorresponding to the solidity curves of FIG. 5 and where a C-D diffuser101 is defined for less than half of the span positions of the channel96. One prior art curve (“PRIOR ART”) is illustrated as well as severalexample curves 108, 109, 110. As illustrated by example curves 108, 109,110, flow through the channel 96 corresponds to a leading edge relativemach number less than or equal to about 0.8 Mach at cruise for each spanposition of the C-D diffuser 101.

The prior art curve corresponds to a C-D diffuser 101 extending from a0% span position to about 16% span position at about 0.78 Mach. Theexample curve 108 corresponds to a C-D diffuser 101 extending from abouta 0% span position to about a 14% span position at about 0.78 Mach. Theexample curve 109 corresponds to a C-D diffuser 101 extending from abouta 0% span position to about a 13% span position at about 0.71 Mach. Theexample curve 110 corresponds to a C-D diffuser 101 extending from abouta 0% span position to about a 15% span position at about 0.73 Mach. Therelatively low solidity arrangements corresponding to example curves108, 109, 110 further reduces engine weight, thereby further reducingfuel consumption.

FIG. 8 illustrates a relationship between airfoil stagger angle (α) andthe average span (AVERAGE SPAN %) for various arrangements of blades 74corresponding to the solidity curves of FIG. 5 and the relative Machnumber curves of FIG. 7. The stagger angle (α) values in FIG. 8 aregiven relative to an axial direction, or plane parallel to the enginelongitudinal axis A. One prior art curve (“PRIOR ART”) is illustrated aswell as several example curves 108, 109, 110.

The prior art curve corresponds to a blade having an airfoil staggerangle (α) of about 18 degrees at about 16% span position. The examplecurve 208 corresponds to a blade 74 having an airfoil stagger angle (α)of about 12 degrees at about a 14% span position. The example curve 209corresponds to a blade 74 having an airfoil stagger angle (α) of about11 degrees at about a 13% span position. The example curve 210corresponds to a blade 74 having an airfoil stagger angle (α) of about13 degrees at about a 15% span. In some examples, the airfoil staggerangles (α) of the curves shown in FIG. 8 at these given span positionscorrespond to a location where the diffuser 101 transitions from aconverging-diverging configuration to a divergent without convergingarrangement as discussed above.

Other airfoil stagger angles (α) are contemplated. In some examples, theC-D diffuser 101 corresponds to an airfoil stagger angle (α) equal to orless than about 15 degrees along each span position. In other examples,the C-D diffuser 101 corresponds to an airfoil stagger angle (α) equalto or less than about 10 degrees along each span position. In anotherexample, the C-D diffuser 101 corresponds to an airfoil stagger angle(α) less than or equal to about 15 degrees, and greater than or equal toabout 10 degrees, for each span position. In some examples, the airfoilstagger angle (α) is less than or equal to about 60 degrees for eachspan position. In one example, the airfoil stagger angle (α) is lessthan or equal to about 55 degrees for each span position. It should beappreciated that the various airfoil stagger angles (α) can be utilizedwith any of the solidities and relative Mach numbers disclosed herein toimprove airfoil efficiency.

Engines made with the disclosed architecture, and including fan sectionarrangements as set forth in this application, and with modificationscoming from the scope of the claims in this application, thus providevery high efficient operation, relatively high stall margins, and arecompact and lightweight relative to their thrust capability. Two-spooland three-spool direct drive engine architectures can also benefit fromthe teachings herein.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

While this invention has been disclosed with reference to oneembodiment, it should be understood that certain modifications wouldcome within the scope of this invention. For that reason, the followingclaims should be studied to determine the true scope and content of thisinvention.

What is claimed is:
 1. A gas turbine engine comprising: a propulsorincluding a rotor hub rotatable about an engine longitudinal axis and anarray of blades circumferentially spaced about the rotor hub, and thearray of blades including 18 or fewer blades; a geared architecture; acompressor section including a low pressure compressor and a highpressure compressor; a turbine section including a low pressure turbineand a high pressure turbine, wherein the high pressure turbine drivesthe high pressure compressor, and the low pressure turbine drives thepropulsor through the geared architecture; a combustor section arrangedbetween the compressor section and the turbine section; wherein eachblade of the array of blades includes pressure and suction sides andextends in a radial direction from a 0% span position at an inner flowpath location to a 100% span position at an airfoil tip, adjacent bladesin the array of blades including a first blade and a second blade, afacing pressure side of the first blade and a facing suction side of thesecond blade defining a channel having a width that varies in achordwise direction between the facing pressure and suction sides at agiven span position of the first and second blades; wherein the width ateach pressure side location of the first blade along the channel isdefined as a minimum distance from the respective pressure side locationto a location along the suction side of the second blade, the width ofthe channel converges in the chordwise direction to establish a throatfor at least some span positions greater than 5% span and less than halfof the span positions, the throat is established at a location spaced adistance from an inlet of the channel, the distance is greater than aradius defined by a leading edge of the first blade at the same spanposition, and the inlet extends along the width of the channelassociated with the leading edge of the first blade; and wherein thearray of blades has a solidity defined by a ratio of an airfoil chordover a circumferential pitch, and a ratio of the width to the solidityat each span position in which the channel converges is between 0.50 and0.85.
 2. The gas turbine engine as set forth in claim 1, wherein thepropulsor section includes a pressure ratio of less than or equal to1.45 across the blades alone at cruise at 0.8 Mach and 35,000 feet. 3.The gas turbine engine as set forth in claim 2, wherein the gearedarchitecture is an epicyclic gear train, and each of the low pressureand high pressure compressors and the low pressure and high pressureturbines includes a plurality of stages.
 4. The gas turbine engine asset forth in claim 3, wherein the geared architecture defines a gearreduction ratio of greater than 2.6.
 5. The gas turbine engine as setforth in claim 4, wherein a low corrected tip speed of the array ofblades is less than 1150 feet per second.
 6. The gas turbine engine asset forth in claim 5, wherein the low pressure turbine drives the lowpressure compressor and an input of the geared architecture.
 7. The gasturbine engine as set forth in claim 6, wherein the low pressure turbineincludes an inlet, an outlet, and a low pressure turbine pressure ratiogreater than 5, and the low pressure turbine pressure ratio is a ratioof a pressure measured prior to the inlet as related to a pressure atthe outlet prior to any exhaust nozzle.
 8. The gas turbine engine as setforth in claim 7, wherein the solidity is between 1.6 and 2.5 for eachof the span positions in which the channel converges.
 9. The gas turbineengine as set forth in claim 8, wherein a stagger angle of each blade ofthe array of blades relative to the engine longitudinal axis is lessthan or equal to 15 degrees at span positions in which the channelconverges and diverges.
 10. The gas turbine engine as set forth in claim9, wherein the solidity is less than or equal to 2.5 at 0% span.
 11. Thegas turbine engine as set forth in claim 10, wherein the epicyclic geartrain is a star gear system.
 12. The gas turbine engine as set forth inclaim 11, wherein the solidity is between 2.0 and 2.5 for each of thespan positions in which the channel converges.
 13. The gas turbineengine as set forth in claim 10, wherein the epicyclic gear train is aplanetary gear system.
 14. The gas turbine engine as set forth in claim13, wherein the solidity is between 2.0 and 2.5 for each of the spanpositions in which the channel converges.
 15. The gas turbine engine asset forth in claim 14, wherein each blade of the array of bladescomprises a composite material.
 16. The gas turbine engine as set forthin claim 8, wherein the solidity is less than or equal to 2.5 at 0%span.
 17. The gas turbine engine as set forth in claim 16, wherein thechannel converges and diverges at span positions greater than 20% span.18. The gas turbine engine as set forth in claim 16, wherein the ratioof the width to the solidity at each span position is greater than orequal to 0.50.
 19. The gas turbine engine as set forth in claim 16,wherein the epicyclic gear train is a planetary gear system.
 20. The gasturbine engine as set forth in claim 19, wherein the solidity is greaterthan or equal to 1.0 at 0% span.
 21. The gas turbine engine as set forthin claim 20, wherein the ratio of the width to the solidity at each spanposition is greater than or equal to 0.70.
 22. The gas turbine engine asset forth in claim 20, wherein a stagger angle of each blade of thearray of blades relative to the engine longitudinal axis is less than orequal to 15 degrees at span positions in which the channel converges anddiverges.
 23. The gas turbine engine as set forth in claim 22, whereinthe channel converges and diverges at span positions greater than 10%span.
 24. The gas turbine engine as set forth in claim 23, wherein theratio of the width to the solidity at each span position is greater thanor equal to 0.70.
 25. The gas turbine engine as set forth in claim 24,wherein: the channel diverges without converging at span positions from100% span to less than or equal to 90% span; and each blade of the arrayof blades comprises a composite material.
 26. The gas turbine engine asset forth in claim 20, wherein the turbine section includes amid-turbine frame between the high pressure turbine and the low pressureturbine, the mid-turbine frame supports bearing systems in the turbinesection, and the mid-turbine frame includes airfoils in a core flowpath.
 27. The gas turbine engine as set forth in claim 26, wherein thechannel converges and diverges at span positions greater than 16% span.28. The gas turbine engine as set forth in claim 27, wherein a staggerangle of each blade of the array of blades relative to the enginelongitudinal axis is less than or equal to 15 degrees at span positionsin which the channel converges and diverges.
 29. The gas turbine engineas set forth in claim 28, wherein the solidity is between 2.0 and 2.5for each of the span positions in which the channel converges.
 30. Thegas turbine engine as set forth in claim 29, wherein: each blade of thearray of blades comprises a composite material; the ratio of the widthto the solidity at each span position is greater than or equal to 0.70;the channel diverges without converging at span positions from 100% spanto less than or equal to 70% span; and flow through the channel at spanpositions where the channel converges and diverges corresponds to aleading edge relative mach number less than or equal to 0.8 Mach atcruise.